Differential pitch-control to optimize co-rotating stacked rotor performance

ABSTRACT

A rotor system for a rotorcraft having at least one rotor blade pair operably associated with a differential pitch assembly operably for controlling a pitch angle of an upper rotor blade and the lower rotor blade in the rotor blade pair. Operation of the differential pitch assembly changes the pitch of the upper rotor blade more severely than the pitch of the lower rotor blade. As such, the rotor system is configured to provide optimum pitch of the upper and lower rotor blades during a helicopter mode and an airplane mode.

TECHNICAL FIELD

The present application relates in general to the field of rotor systems for rotorcraft.

DESCRIPTION OF THE PRIOR ART

There are many different types of rotorcraft, including helicopters, tandem rotor helicopters, tiltrotor aircraft, four-rotor tiltrotor aircraft, tilt wing aircraft, and tail sitter aircraft. In all of these rotorcraft, thrust and/or lift is generated by air flowing through a rotor disk formed by a plurality of rotating rotor blades. The plurality of rotor blades are mechanically coupled with and substantially evenly spaced about a rotatable mast, which provides rotational motion to the plurality of rotor blades. Each of the plurality of rotor blades is independently rotatable to affect a pitch of the blade. Varying the pitch of the plurality of blades affects lift and the direction of thrust produced by the rotating plurality of blades.

FIG. 1 depicts a military tiltrotor aircraft 101 with conventional rotor hubs 107 a and 107 b. Rotor hubs 107 a and 107 b are mechanically coupled to nacelles 103 a and 103 b, respectively. Nacelles 103 a and 103 b are rotably attached to wing members 105 a and 105 b, respectively. Wing members 105 a and 105 b are rigidly fixed to fuselage 109. Rotor hubs 107 a and 107 b have a plurality of rotor blades 111 a and 111 b, respectively. The tiltrotor aircraft 101 of FIG. 1 is depicted in helicopter mode, with nacelles 103 a and 103 b directed up.

FIG. 2 depicts a commercial tiltrotor aircraft 201 with conventional rotor hubs 207 a and 207 b. Rotor hubs 207 a and 207 b are mechanically coupled to nacelles 203 a and 203 b, respectively. Nacelles 203 a and 203 b are rotably attached to wing members 205 a and 205 b, respectively. Wing members 205 a and 205 b are rigidly fixed to fuselage 209. Rotor hubs 207 a and 207 b have a plurality of rotor blades 211 a and 211 b, respectively. FIG. 2 depicts tiltrotor aircraft 201 in airplane mode, with nacelles 203 a and 203 b directed forward.

It is often desirable to utilize a multiple rotor disks in a stacked configuration, the multiple rotor disks rotating about the same axis of rotation, to increase the lift and/or thrust of a rotorcraft. In operation, each rotor disk may experience a different aerodynamic condition, depending on variables such as forward speed and orientation of the rotor hub. There is a need for rotor system which can variably adjust the pitch of the rotor blades of each rotor disk in a simple and efficient means.

There are many rotorcraft rotor systems well known in the art; however, considerable room for improvement remains.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the system of the present application are set forth in the appended claims. However, the system itself, as well as, a preferred mode of use, and further objectives and advantages thereof, will best be understood by reference to the following detailed description when read in conjunction with the accompanying drawings, wherein:

FIG. 1 is a perspective view of a prior art tiltrotor aircraft in helicopter mode;

FIG. 2 is a front view of the prior art tiltrotor aircraft of FIG. 1, in airplane mode;

FIG. 3 is a front view of a tiltrotor aircraft, in airplane mode, having a rotor system according the preferred embodiment of the present application;

FIG. 4 is a perspective view of the rotor system from the aircraft in FIG. 3, according to the preferred embodiment of the present application;

FIG. 5 is a close-up perspective view of the rotor system from the aircraft in FIG. 3, according to the preferred embodiment of the present application;

FIG. 6 is a stylized schematic view of the rotor system from the aircraft in FIG. 3, while in airplane mode, according to the preferred embodiment of the present application; and

FIG. 7 is a stylized schematic view of the rotor system from the aircraft in FIG. 3, while in helicopter mode, according to the preferred embodiment of the present application.

While the system of the present application is susceptible to various modifications and alternative forms, specific embodiments thereof have been shown by way of example in the drawings and are herein described in detail. It should be understood, however, that the description herein of specific embodiments is not intended to limit the invention to the particular forms disclosed, but on the contrary, the intention is to cover all modifications, equivalents, and alternatives falling within the spirit and scope of the system of the present application as defined by the appended claims.

DESCRIPTION OF THE PREFERRED EMBODIMENT

Illustrative embodiments of system of the present application are described below. In the interest of clarity, not all features of an actual implementation are described in this specification. It will of course be appreciated that in the development of any such actual embodiment, numerous implementation-specific decisions must be made to achieve the developer's specific goals, such as compliance with system-related and business-related constraints, which will vary from one implementation to another. Moreover, it will be appreciated that such a development effort might be complex and time-consuming but would nevertheless be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure.

In the specification, reference may be made to the spatial relationships between various components and to the spatial orientation of various aspects of components as the devices are depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of the present application, the devices, members, apparatuses, etc. described herein may be positioned in any desired orientation. Thus, the use of terms such as “above,” “below,” “upper,” “lower,” or other like terms to describe a spatial relationship between various components or to describe the spatial orientation of aspects of such components should be understood to describe a relative relationship between the components or a spatial orientation of aspects of such components, respectively, as the device described herein may be oriented in any desired direction.

The present application represents a rotor system for a rotorcraft and a rotorcraft incorporating the rotor system. The rotor system includes a rotor hub having a plurality of rotor blade pairs. A differential pitch link assembly is mechanically coupled to each rotor blade pair, for variably controlling a pitch angle of each rotor blade in the rotor blade pair, in tandem. Actuation of swashplate link provides an input to the differential pitch link assembly, thereby changing the pitch of each rotor blade in the rotor blade pair associated with the particular pitch link assembly.

Referring now to FIG. 3 in the drawings, FIG. 3 is a front view of a rotorcraft 311, while in airplane mode. A first rotor hub 317 a is mechanically coupled to a first nacelle 337 a and a second rotor hub 317 b is couple to a second nacelle 337 b. Nacelles 337 a and 337 b are pivotally attached to wing members 333 a and 333 b, respectively. Wing members 333 a and 333 b are rigidly attached to a fuselage 331. Nacelles 337 a and 337 b are configured to pivotally rotate relative to wing members 333 a and 333 b between a helicopter mode, in which nacelles 337 a and 337 b are tilted upward such that rotorcraft 311 flies similar to a conventional helicopter; and an airplane mode in which nacelles 337 a and 337 b are tilted forward such that rotorcraft 311 flies similar to a conventional propeller-driven airplane. FIG. 3 further depicts the preferred embodiment having three rotor blade pairs 349 on each rotor hub 317 a and 317 b.

FIG. 4 in the drawings is a perspective view of rotor hub 317 a from rotorcraft 311, while in helicopter mode. A differential pitch link assembly 301 is coupled to a swashplate 303. An upper rotor yoke 309 and a lower rotor yoke 319 are coupled to and rotatable by a mast 245 about axis of rotation 335. Pitch settings for each rotor blade pair 349 are controlled independently. As discussed herein, pitch settings of rotor blades 323 a and 323 b are controlled in tandem for each rotor blade pair 349. Axis of rotation 335 of mast 245 is located at an intersection rotor blade pitch axes 321, shown in FIG. 5. While the preferred embodiment depicts three rotor blade pairs 349 in each rotor hub 317 a and 317 b, the scope of the present invention is not so limited. Rather, any suitable plurality of rotor blade pairs 349 may be implemented, and the present application contemplates such alternative embodiments.

Now referring to FIG. 5, a more detailed perspective view of rotor hub 317 a is illustrated. An upper blade pitch control member 343 a is used, in part, to couple upper rotor yoke 309 to upper rotor blade 323 a. Similarly, a lower blade pitch control member 343 b is used, in part, to couple lower rotor yoke 319 to lower rotor blade 323 b. For clarity, FIG. 5 does not illustrate upper rotor blade 323 a and lower rotor blade 323 b.

Rotor hub 317 a has at least one differential pitch link assembly 301 associated with at least one rotor blade pair 349, used to differentially manipulate the pitch of upper rotor blade 323 a and lower rotor blade 323 b, according to optimum pitch angles. Differential pitch link assembly 301 comprises a swashplate link 325 operably coupled to a lower pitch horn 339, and an upper pitch horn 313 operably coupled to lower pitch horn 339 via an intermediate link 315. Upper pitch horn 313 is operably associated with upper blade pitch control member 343 a. Similarly, lower pitch horn 339 is operably associated with lower blade pitch control member 343 b. Swashplate 303 is configured to selectively actuate each swashplate link 325 in a swashplate actuation direction 305. In the preferred embodiment, three differential pitch link assemblies 301 are used to control three rotor blade pairs 349; however, it should be appreciated that greater or fewer differential pitch link assemblies 301 may be used in conjunction with a greater or fewer rotor blade pairs 349. Differential pitch link assembly 301 is applicable to rotor systems having an upper rotor disk assembly 307 and a lower rotor disk assembly 341, each disk assembly 307 and 341 configured to be co-axial and co-rotating about an axis of rotation 335.

In the preferred embodiment, upper rotor disk assembly 307 is comprised of three upper blade pitch control members 343 a coupled to upper rotor yoke 309. Similarly lower rotor disk assembly 341 is comprised of three lower blade pitch control members 343 b coupled to lower rotor yoke 319. Each differential pitch link assembly 301, swashplate 303, upper rotor disk assembly 307, and lower rotor disk assembly 341, are rotatable by a mast 345 about axis of rotation 335.

FIG. 6 represents a stylized schematic view of rotor hub 317 a in airplane mode. During airplane mode, the upper rotor disk assembly 307 experiences relatively the same air inflow velocity as lower rotor disk assembly 341. Therefore, airplane mode upper air flow 329 a and airplane mode lower air flow 329 b are approximately the same. As such, during airplane mode, upper rotor disk collective pitch angle D1 is substantially similar to lower rotor disk collective pitch angle C1.

FIG. 7 represents a stylized schematic view of rotor hub 317 a in helicopter mode. During helicopter mode, upper rotor disk assembly 307 accelerates air towards lower rotor disk assembly 341. As such, lower rotor disk assembly 341 experiences a higher air inflow velocity than experienced by upper rotor disk assembly 307. Helicopter mode lower air flow 327 b is significantly higher than helicopter mode upper air flow 327 a. In order to provide optimal performance, a lower disk collective pitch angle A1 is higher than an upper disk collective pitch angle B1 so as to accommodate the difference the velocities of air flow 327 a and air flow 327 b. In order to achieve optimal disk angles while using the same swashplate link 325 to control both upper and lower pitch horns 313 and 339; an operable pitch horn length L1 is longer than an operable pitch horn length L2. Because operable pitch horn length L1 is longer than L2, actuation of swashplate link 325 changes the upper rotor disk collective pitch angle B1 more severely than lower rotor disk collective pitch angle A1. When converting from airplane mode to helicopter mode, swashplate link 325 is actuated up towards upper and lower rotor disk assemblies 307 and 341.

Upper blade pitch control member 343 a is used, in part, to couple upper rotor yoke 309 to upper rotor blade 323 a. Similarly, lower blade pitch control member 343 b is used, in part, to couple lower rotor yoke 319 to lower rotor blade 323 b. Each of rotor blade pair 349 comprises upper rotor blade 323 a and lower rotor blade 323 b. Each rotor blade 323 a and 323 b comprises an airfoil shape capable of producing lift when air is moved over the airfoil shape. As such, the amount of lift can be controlled by adjusting the pitch of upper rotor blade 323 a and lower rotor blade 323 b, in tandem with an associated differential pitch link assembly 301. Upper rotor yoke 309 and lower rotor yoke 319 are located in two different planes along axis of rotation 335. As such, upper rotor blade 323 a and lower rotor blade 323 b operate approximately in the separate spatial planes as defined by the location of upper rotor yoke 309 and lower rotor yoke 319, respectively. Upper rotor yoke 309 and lower rotor yoke 319 are clocked by an angle A, about axis of rotation 335; as such, upper rotor blade 323 a and lower rotor blade 323 b are positioned about the same angle A. In the preferred embodiment, Angle A is approximately 30 degrees; however, Angle A may also be other angles depending on design factors; such as, number of rotor blade pairs 349 and size of rotor blades 323 a and 323 b.

FIG. 3 depicts the preferred embodiment of a rotorcraft 311 incorporating two rotor hubs 317 a and 317 b, and a plurality of rotor blade pairs 349, operably associated with each rotor hub 317 a and 317 b. Masts 345 (one shown in FIG. 4) extend from a transmission (not shown) within each nacelle 337 a and 337 b of rotorcraft 311. It should be noted that, while rotorcraft 311 is depicted in FIG. 3 as being a tiltrotor aircraft, the scope of the present invention is not so limited. Rather, the system of the present application contemplates rotorcraft 311 being any type of rotorcraft, such as a helicopter, a tandem rotor helicopter, a tiltrotor aircraft, a four-rotor tiltrotor aircraft, a tilt wing aircraft, or a tail sitter aircraft.

The system of the present application provides significant advantages, including: (1) providing a way to efficiently control the pitch of multiple rotor blades according to optimum pitch angles during a helicopter mode and an airplane mode; (2) providing a way to utilize a plurality rotor blades in a rotorcraft without the added control mechanism weight found in conventional rotorcraft; (3) providing a way to utilize a plurality rotor blades in a rotorcraft while keeping the rotor hub size as small as possible.

It is apparent that a system with significant advantages has been described and illustrated. Although the system of the present application is shown in a limited number of forms, it is not limited to just these forms, but is amenable to various changes and modifications without departing from the spirit thereof. 

The invention claimed is:
 1. A rotor system for a rotorcraft, comprising: a rotor mast; an upper rotor blade and a lower rotor blade, both the upper rotor blade and the lower rotor blade being coupled to the rotor mast so as to rotate about a singular axis of rotation and in a singular direction; and a differential pitch assembly operably associated with the upper rotor blade and the lower rotor blade, so as to control a pitch angle of both the upper rotor blade and the lower rotor blade in tandem; wherein the differential pitch assembly is configured to establish the pitch angle of the upper rotor blade approximately similar to the pitch angle of the lower rotor blade when the rotorcraft is in an airplane mode; wherein the differential pitch assembly is configured to establish the pitch angles of the upper rotor blade and the lower rotor blade such that so that the pitch angle of the upper rotor blade is smaller than the pitch angle of the lower rotor blade when the rotorcraft is in a helicopter mode.
 2. The rotor system according to claim 1, wherein the lower rotor blade is clocked forward of the upper rotor blade by approximately 30 degrees about the axis of rotation.
 3. The rotor system according to claim 1, wherein the helicopter mode entails the lower rotor blade experiencing a higher inflow air velocity than experienced by the upper rotor blade.
 4. The rotor system according to claim 1, wherein the airplane mode entails the lower rotor blade experiencing a relatively similar inflow air velocity to that as experienced by the upper rotor blade. 